1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with tip cooling and sealing for an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
The performance of a highly loaded turbine is a strong function of the turbine blade tip clearance and leakage flow. A large running tip clearance with high leakage flow will induce high performance losses. Therefore, blade tip section sealing and leakage flow reduction and tip clearance gap should be addressed as a single problem. A prior art turbine blade includes a squealer tip rail that extends around a perimeter of the airfoil flush with the airfoil walls on the pressure and suction sides and forms an inner squealer pocket. The main purpose for using a squealer tip in the blade design is to reduce the blade tip leakage and to provide for rubbing capability of the blade tip with a stationary shroud surface in the turbine. FIG. 3 shows a turbine blade with a squealer tip with a secondary hot gas flow migration pattern over the blade tip section. The hot gas leakage flow flows across the blade tip section from the pressure side over the squealer pocket to the suction side. This hot gas migration from the blade lower span height is due to a pressure gradient formed between the pressure side and the suction side. Hot gas 25 flows in to the pocket, the secondary leakage flow 26 flows over the pocket, and a vortex flow 27 is formed next to the suction side tip rail. Smaller vortices are also formed that pass over the two rows of cooling air holes that open into the pocket and extend along the tip rails.
FIG. 4 shows a pressure distribution along the blade periphery at the tip location. A pressure level on the pressure side (P/S) is higher than on the suction side (S/S) of the blade. The highest pressure is found at the blade leading edge. A pressure differential 33 is formed between the two sides of the blade and varies along the chordwise length of the blade.
FIG. 5 shows a temperature profile of the hot gas flow relative to the blade. A temperature level at the blade tip section is lower than a temperature level at the blade lower span or root. The highest gas flow temperature 34 is located at a mid-chord region where the hot gas flows over the blade tip. Because of this, the hot gas flow at the blade tip section can be used for cooling at the blade aft section.